Enthalpy

Solar Thermal Rocket

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Dear visionary inventors, megalomaniac engineers and audacious explorers, biggrin.png

 

Chemical rocket engines aren't up to our desire to hop in the Solar system: go quickly to Mars, deviate Earth-threatening objects, and so many more missions. We need a higher ejection speed to save propellant mass, but this takes more energy than chemical reactions bring. One possibility is to tap Sunlight instead of transporting the energy.

I suggest - as others did - to directly heat the propellant with Sunlight. Converting first to electricity would enable even higher ejection speeds, but direct heating is energy-efficient, so for a long weak thrust, the collector area is feasible - smaller and of cheaper materials than Solar cells or even a heat sink.

Material sublimation limits thermal designs to <3000K, so only hydrogen improves the ejection speed over a combustion. My plan is to also dissociate a part of this hydrogen to increase the ejection speed further.

--------------

The heater that catches concentrated Sunlight and transfers heat to hydrogen is of tungsten alloy. At 2800K =2527°C, sublimation thins it by 45µm in 14 days, from Plansee's doc.

post-53915-0-45137500-1371757299.png

30mbar in the heating chamber let 23% of injected H2 split to atomic H* for performance. Expansion to 1Pa in the nozzle leaves 15.6MJ/kg from 92.8MJ/kg in the chamber, for isp = 1267s = 12.4km/s. tongue.png

Mean 2800K is already a lot, as sublimation is very sensitive to it. The nozzle can't grow much because the molecules' mean free path is already ~10mm. A lower chamber pressure dissociates more hydrogen and improves the ejection speed, but needs much more heating power, as the nozzle gets inefficient - recombining hydrogen is a hard task. But more pressure brings during some flight sequences more thrust traded against ejection speed: for instance, 0.8bar and 2093K from the same Sunlight concentrator and more hydrogen throughput push *2.1 times stronger with isp=800s.

 

More to come. A former version, partially inaccurate, began there

http://saposjoint.net/Forum/viewtopic.php?f=66&t=2164

and an actual version has begun there

http://saposjoint.net/Forum/viewtopic.php?f=66&t=2164&p=42029#p42029

which I plan to describe more concisely in the coming few days on ScienceForums, so stay tuned!

Marc Schaefer, aka Enthalpy

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for long distance travel or interstellar travel,

 

the problem is not rockets and fuel,

but the human element,

mainly how to stop individuals from going insane.

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How do you connect the reflector to the engine and inhabited modules? A solar sail which your diagram resembles drags the module behind it so tensioned carbon fiber would light and strong but pushing that mirror seems to be a significant problem...

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Much Sunlight power is needed, so most uses will have several concentrators as big as the launcher's fairing permits. They can consist of metallized graphite honeycomb, like usual satellite antennas. One chamber per concentrator gives redundancy and eases the orientation and the ground tests.

D=4.4m provides near Earth 16.6kW if 80% are used, enough for 195mg/s hydrogen to push 2.4N.

 

post-53915-0-02687000-1371763411.png

 

The chamber's light inlet must be small because it's hot: at D=38mm, blackbody's radiation loses 19% of the incoming power. Hence the chamber must be close enough to the primary mirror (here 0.92 diameter) that the Sun's image fits through the inlet.

The small secondary mirror makes light more parallel. At D=0.2m, the temperature can be kept bearable by the reflective surface.

The nozzle must be oriented, using the secondary mirror and possibly more.

A different telescope formula can bring advantages, but it must keep light's path short.

Marc Schaefer, aka Enthalpy

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Oops.wacko.png

 

I wanted nearly parallel light converging to a small image of the Sun, and this is not compatible. A small image implies a strong convergence. Thermo's second law tells it, but knowing geometrical optics would have been even better...doh.gif

 

The good news is that a small image accepts a long path in the telescope. Just one example here, where rotating mirrors orient the thrust, and this movement adds no aberration:

 

post-53915-0-76040200-1371821125.png

 

Stacking many mirrors in a launcher's fairing constraints such designs. And with several engines and a craft, the control must avoid collisions, as the jet impact must be hot and corrosive.

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Light puts pressure on the mirror, which your rocket must overcome, reducing efficiency.

 

If you make both your primary and secondary mirrors with holes in the middle, you may be able to concentrate light on a circle on the outside of the engine and direct its exhaust through the mirror holes. Thus, combining light pressure on the primary mirror and engine exhaust to increase efficiency. The secondary mirror may be larger and more massive, which may offset the added efficiency. The mirror shapes are more difficult to design and may to difficult to build. My vision is a very short focal length with only two mirrors. Perhaps clever engineering can reduce total mass to near that of a three mirror system. The mirror math is beyond my capability.

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This is how the chamber could be made, I believe - new developments always bring surprises. Hot parts are of tungsten alloy. Spark machining shall make the many thin deep shapes, and an electron beam weld the parts. Click to magnify the image.


post-53915-0-85613700-1372013074_thumb.png

and the front view:

post-53915-0-11384700-1372013090.png


The Sun's virtual image (D=38mm, 20kW and 18MW/m2 from 4.4m concentrators) enters the heater's cavity (40mm) with 1:1.85 convergence. The heater absorbs light and transfers heat by conduction to hydrogen in channels. 63µK*m2/W through tungsten and <37µK*m2/W through hydrogen drop only 130K at the heater's beginning, with ~1.3MW/m2 absorbed there and 60% reflected. At depths where direct and reflected light have dropped, helical fins in the cavity increase the absorption, and a converging cone as well. The bottom finishes gently to heat the hydrogen with little drop.

The heater's outer stern would radiate much (2800K and e~0.32) but the regenerator's inner face reflects it (a=0.02). The absorbed fraction (~1500W) pre-heat hydrogen by 520K.

A ruminator surrounds the incoming light cone to absorb much of the light emitted through the cavity's inlet and transfer it to hydrogen (+1060K): of 4380W from a 40mm blackbody, an infinite cone would catch 3320W, but 340W less if truncated to D=200+38mm (sketched shorter).

The nozzle's throat (22mm) and upper divergent emit >1320W light; cooling regeneratively the lower divergent would save much of it.

The regenerator holds the chamber and nozzle by the less hot hydrogen inlet, and the regenerator itself holds by its cool hydrogen feed. Wall conduction (~800W at the cavity's mouth) thus pre-heats the hydrogen.

Marc Schaefer, aka Enthalpy

Edited by Enthalpy

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Light puts pressure on the mirror, which your rocket must overcome, reducing efficiency.

 

If you make both your primary and secondary mirrors with holes in the middle, you may be able to concentrate light on a circle on the outside of the engine and direct its exhaust through the mirror holes. Thus, combining light pressure on the primary mirror and engine exhaust to increase efficiency. The secondary mirror may be larger and more massive, which may offset the added efficiency. The mirror shapes are more difficult to design and may to difficult to build. My vision is a very short focal length with only two mirrors. Perhaps clever engineering can reduce total mass to near that of a three mirror system. The mirror math is beyond my capability.

On second thought, only a primary mirror is necessary. But, holding shape of mirror is a challenge.

Edited by EdEarl

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My answers to some points evoked here - sorry to be late...

 

Interstellar travel does not exist, not even as a proof-of-concept, because we have no propulsion for it.

 

A manned travel to Mars for instance would take under 100 days. Take submarine crews for that, they usually don't get crazy when confined over this lapse.

 

Radiation pressure is 4.5µPa near Earth. This is 68µN on a 4.4m dish. It's the very reason why we don't use it every day for space transport. 2N by my Solar thermal rocket is unconveniently little, but usable for some missions. That said, I do like Solar sails, but we still need comfortable solutions for hectares and square kilometers of sail.

 

I will not detail how to unfold and control the concentrators. This is not new technology; spacecraft designers do it better than I, for cases already much more difficult than the engine I describe.

 

Multi-meter telescopes in space make pictures limited by diffraction, far below the micron. A concentrator that focusses Sunlight to make heat is nowhere as difficult. Its accuracy compares rather with a radio dish antenna, both in shape and pointing, for which a standard carbon honeycomb is perfect.

 

The thrust must be arbitrarily orientable versus Sunlight. From Earth to Mars for instance, the craft would accelerate perpendicularly to the Sun's direction, but decelerate strongly Sunwards. Other missions have still other needs. In short: steer in all directions. If it simplifies other parts, especially the chamber, an added mirror (1-1.5% light loss, little mass, limited cost) is perfectly welcome. Steering without excessive optical aberration is nothing obvious, and I'm happy that my three mirrors show it's feasible, though optics designers will find better combinations.

 

Any spacecraft will cumulate several concentrators (a reasonable number for some missions and designs) so the craft will control its attitude from independent orientation of individual engines.

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To bring a payload from Low-Earth- (Leo) to a Geosynchronous Orbit (Gso), a Solar thermal rocket would use its small push all the way to save time, but this takes more performance than the usual very elliptical transfer (Gto).

From an equatorial Leo (sea launch, Alcantara...) the required Delta-V is the speed difference between the low and the high orbit, if I didn't botch it. Starting from 7675m/s at 6366+400km to avoid drag, the transfer stage must add 4600m/s to achieve the 3075m/s at 6366+35798km. This costs 741m/s more than the Hohmann Gto.

From other launch sites, the orbit's inclination must be cancelled. The very elliptic Gto does it at apogee where speed is small, but the continuous push lacks this possibility. Fortunately, isp=1267s absorbs this as well.

Kourou reaches 5.2° inclination at Leo. To estimate the added cost, I keep the inclination over the first 2600m/s, and compensate it over the last 2000m/s needed for altitude; the side speed to cancel is 462m/s (begin) to 278m/s (if done at Gso), so I take mean 370m/s. The engines shall push flat for 2*90° per orbit, and for 2*90°, tilted as 1,11*370m/s * sin = 1000m/s (1,11 because the tilted push extends 45+45° from the optimum point). When tilted, only 91% of the thrust is equatorial, so these 1000m/s cost 1097m/s performance, and the overhead nears 97m/s, totalling 4697m/s to Gso. A Gto would waste 15m/s from Kourou's latitude.

Cape Canaveral reaches 28.5° (and Tanegashima 30°), giving 3779m/s side speed at 400km and 1505m/s at height, for mean 2642m/s. Here I tilt the thrust as a cosine function over one orbit. At maximum tilt, the relative components of thrust are S and C, kept for all orbits of the transfer, which is not optimum. Only 0.5*S and nearly (1+C)/2 act as a mean, so side 2642m/s versus 4400m/s along the path leave 73.5% efficiency: the total cost nears 5986m/s and the overhead 1586m/s. This wastes even more performance than de-inclining at final altitude but saves time.

The optimum is obviously elsewhere - someone with a liking for it shall investigate. I take 5700m/s.

 

=================================================================

 

A Falcon-9 shall illustrate the transfer from Leo to Gso - click to view the drawing's full splendor. Starting from Cape Canaveral, the launcher puts 10,0t on a 28,5° 400km orbit. The Falcon may need reinforcement for the longer fairing.

 

post-53915-0-66679200-1372700983_thumb.png

 

Minus the adapter, the transfer stage starts with 9350kg. To provide 5700m/s as estimated previously, it ejects 3440kg hydrogen at isp=1267s. 20 days thrust (plus some 5 days eclipse) take 10 Solar thermal engines with 4.6m concentrators. Not that huge.

The balloon tank of thin steel, foam and multilayer insulation weighs 143kg. Polymer straps hold it to a truss of welded AA7020 weighing 205kg that links to the launcher and the payload. A cryocooler keeps the hydrogen liquid.

Each engine weighs 100kg, of which 50kg are the concentrator and 20kg the chamber, using nickel or niobium rather than tungsten where possible.

300kg of sensors, datacomms, control and unaccounted items leave a payload of 4262kg in Gso. That's roughly twice the capacity of chemical stages.

-----

Here the Solar thermal stage is expended at each launch and continues to a park orbit. It can also come back to Leo, what ESA calls a "tugboat":

  • Reusing it saves launch mass, even though the way back needs some hydrogen;
  • It can bring a satellite down to Leo for repair;
  • A lighter launch mission can bring extra hydrogen for a following overweight mission;
  • The launcher can bring the hydrogen and the satellite separately. This puts in Gso the full launcher's Leo capacity;
  • A flexible long-range vehicle between Leo and Gso can bring resupply or remote repairs to several satellites, push aside the lost ones...

Marc Schaefer, aka Enthalpy

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Escaping Earth or an other body is best done with one or few kicks at the perigee. The Solar thermal engine is bad here: its faint thrust used briefly at each orbit would take too long and demand too many concentrators. Escaping by continuous rise of a circular orbit is inefficient, letting a Solar thermal engine use as much propellant as an oxygen-hydrogen engine. Chemical engines shall escape planets.

Providing little more speed at perigee than the escape minimum leaves much speed after escape because energy goes as speed squared
http://en.wikipedia.org/wiki/Oberth_effect
so I checked for a high-energy mission how the chemical and Solar engines shall share optimally the provision of speed beyond escape.

From 200km above Earth, where speed is 7791m/s and needs 3227m/s more to escape:

  • An oxygen and hydrogen engine with isp=465s best brings 4054m/s = 3227 + 827m/s, leaving 4346m/s beyond escape;
  • An oxygen and dense fuel engine with isp=395s best brings 3789m/s = 3227 + 562m/s, leaving 3561m/s beyond escape.

The wide optimum adapts to practical arguments. The needed performance nears a Mars transfer or a direct injection in Gso.

Marc Schaefer, aka Enthalpy

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The Solar thermal engine's isp=1267s=12424m/s can bring asteroid samples to Earth tongue.png .

 

Take a Falcon-9 for its wide fairing; it puts 10t at 400km Leo. A Leo-to-Gso stage as I describe there
http://www.scienceforums.net/topic/73571-rocket-engine-with-electric-pumps/#entry736092

used as an escape stage injects 2790kg at 4880m/s above Earth's gravity.

 

A Hohmann transfer to 2.58UA in the main asteroid belt takes only 15 months
http://en.wikipedia.org/wiki/Asteroid_belt
it needs additional 1093m/s near Earth and 4684m/s in the belt, while the return leg takes 4684m/s in the belt, and a capsule aerobrakes for free from 12.2km/s. Cumulated 10461m/s would leave 1200kg at reentry eyebrow.gif - but meanwhile samples were taken, and the craft has made manoeuvres.

Braking 4684m/s takes 804kg hydrogen; over 30 days, it needs 7 concentrators of D=4.6m at 2.58UA. These are oversized at 1UA, and the diameter of a secondary or third mirror can limit the power if desired. The chambers for 2.58UA are lighter that previous estimations, and I'm confident that the concentrators can weigh <<3kg/m2, so each engine is maybe 30-50kg. This leaves mass for equipment and samples, all with a single stage.

 

Fascinating: the craft pushing 0.4N per 100kg at 2.58UA can lift off a D=10km asteroid by its Solar engines, and hop from one asteroid to an other, taking samples at each one ohmy.png - maybe capture tiny asteroids provided it sees some with a reasonable speed.

Marc Schaefer, aka Enthalpy

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A mission similar to the main belt asteroids can bring back samples from Jupiter's Trojans:
http://en.wikipedia.org/wiki/Jupiter_Trojan

The main belt return trip leaves room for extras, like sampling both at 2.38 AU and 2.58 AU (+764m/s if you're patient); but Jupiter's orbit needs to care more about speed changes and mass.

The 2-year Hohmann transfer needs 8793m/s over Earth's gravity at perihelion and 5643m/s at aphelion. If again a supplemented Falcon-9 puts 2793kg at 4881m/s, the trip takes 15198m/s. If the parts thrown away only compensated the samples, the craft would arrive at Earth at 822kg.

754, 744 and 472kg hydrogen are ejected; braking over 120 days at 5.2 AU takes 9 engines with 4.6m concentrators, and accelerating there only 6 engines. Concentrators weighing 1kg/m2 seem a reasonable effort, so each engine could weigh 30kg. Tanks of 100µm steel, 20mm foam and 50 layer insulation, hold by polymer straps, weigh 110kg for the first 754+744kg hydrogen and 55kg for the last 472kg - in case the first is thrown away. The sampling drill can stay there as well. That's room for samples, and for the ability to visit several Trojans.

This would leave some 400kg for all craft functions and the re-entry capsule, not including the tanks, engines and 180kg souvenirs. Bonanza! tongue.png

 

Missions to Jupiter use the Oberth effect to brake there, but I suppose the Trojans are too far away. Bigger launchers, like Ariane 5 with the Esc-B, would give a stronger initial kick to a heavier payload, for instance both missions to the main belt asteroids and to the Jupiter Trojans at once.

Marc Schaefer, aka Enthalpy

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Samples from Phobos and Deimos, Mars' moons, are a mission easier than asteroid samples:

http://en.wikipedia.org/wiki/Phobos_(moon)

http://en.wikipedia.org/wiki/Deimos_(moon)

 

From the Hohmann transfer, a supplemented Falcon-9 provides all the 2945m/s above Earth's gravity to 3414kg spacecraft, while Earth's atmosphere makes all the braking on the return leg. The Solar thermal rocket provides Hohmann's 2649m/s near Mars, then 2139m/s to descend to Phobos, some 216m/s to sample a dozen sites there, 789m/s to climb to Deimos, some 216m/s to sample a dozen sites, 1350m/s to climb and escape Mars' gravity, and 2649m/s to come back. This amounts to 10260m/s, easy for the Solar thermal rocket.

The craft weighs 2758kg as is hovers at Phobos. At the moons, the engines operate at 2.1* thrust, or isp=800s, so 2* 12 hops of 1h each at 5mm/s2 comsume as much hydrogen as 684m/s at full isp. This needs only five 4.6m concentrators then. More thrust increase would improve.

The craft weighs still 1494kg when arriving near Earth. Since thick regolith covers the moons, lighter samples should suffice, permitting a smaller launcher and craft.

Marc Schaefer, aka Enthalpy

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Mercury needs 7533m/s then 9611m/s in a Hohmann transfer from Earth, that's why chemical rockets make detours by Venus and Earth and take 7 years with a big launcher.

The Solar thermal rocket goes there in three months. Start as usual with 2793kg at 4881m/s above Earth's gravity:

  • 537kg hydrogen provide the remaining 2652m/s at aphelion, leaving 2256kg. Two 4.6m concentrators take 15 days.
  • 1216kg hydrogen give 9611m/s at perihelion, leaving 1040kg above Mercury's gravity. The two concentrators take 10 days.
  • 112kg hydrogen used close to Mercury give 1245m/s to achieve if desired a low circular orbit (escape: 4250m/s).
    If pushing for 1/7 of the orbital period, where the kicks are about 88% efficient, this takes around one week and leaves 928kg.

Mercury's orbit is eccentric (and tilted). The Hohmann transfer, from Jarret Mathwig's thesis, supposes circular orbits, so a slick mission planner would save performance. Some Oberth effect is also possible; trade isp for thrust there.

Each concentrator may weigh 17kg and the stronger chamber 87kg, so both engines take 208kg. The 28m3 tank with thin steel, foam, multilayer insulation and polymer belts takes 130kg. This leaves 590kg on low Mercury orbit for frame, equipment and instruments. Orbit changes are possible.

Marc Schaefer, aka Enthalpy

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The Pioneer and Voyager probes have seen interesting physics near 90 Astronomical Units. Grass isn't greener there, but it's something like the Termination shock, where our Sun doesn't fully define the medium any more. As well, we could better measure star distances by parallax, and check on the way the Pioneer anomaly and some general physics. Though, the probes were meant and equipped to investigate none of these, and took four decades - so shall we have a new probe there within ten years?

This needs 42780m/s above Sun's gravity (or a bit less), meaning a start at biggrin.png30252m/s over Earth's gravity and own 29785m/s. An Ariane 5 with the Esc-B (public data is incomplete) shall provide 4346m/s over Earth's gravity to 7345kg, oriented as 3199m/s forward and 2941m/s sunwards, so the perihelion is at 0.98 AU and gives 19+15 days to accelerate under 0.99 AU conditions of Sunlight and speed.

A first stage ejects 4756kg hydrogen through fifteen D=4.4m engines to bring 12953m/s. The tank weighs 230kg, a truss 210kg, 11 engines 330kg, leaving 1820kg after separation. More stages would improve.

The second stage ejects 1178kg through the four kept engines, to bring 12953m/s more. 95kg tank, 55kg truss, 120kg engines leave 372kg for the frame, equipment, instruments.

Is a very low perihelion better? The heavier chambers reduced the payload in the case I checked. And if a mission needs a low circular Sun orbit, or a circular polar one, I feel a Solar sail better.

Maybe astronomers would share the Square Kilometer Array, if it's for good science, and if the probe transmits as short bursts, or if a separate frequency wastes no observation time.

Marc Schaefer, aka Enthalpy

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The high specific impulse permits a Jupiter mission that orbits several moons successively and observes in between the planet from varied distances.
http://en.wikipedia.org/wiki/Jupiter
http://en.wikipedia.org/wiki/Moons_of_Jupiter
The Galileo probe had 200kB of Ram and a magnetic tape, so a new design could carry improved instruments.
http://en.wikipedia.org/wiki/Galileo_spacecraft

Atlas V 551 (or Ariane V Me) can put 8900kg on a transfer orbit 1804m/s below geosynchronous, so it shall put 5459kg at 4500m/s above Earth's gravity. The Solar thermal rocket makes the rest, beginning with a 2 years 9 months Hohmann transfer (my mistake at the Trojans).

1596kg hydrogen add 4294m/s to Atlas' 4500m/s and Earth's 29785m/s, leaving 3863kg heading to Jupiter.

1411kg hydrogen accelerate by 5643m/s and leave 2452kg just above Jupiter's gravity well. Twelve D=4.57m engines take 175 days to brake at quasi 5.2AU, lengthening by almost 3 months; their collective consumption there is 93mg/s = 8.04kg/d. Similar concentrators can provide each ~500W electricity as I describe there:

http://saposjoint.net/Forum/viewtopic.php?f=66&t=2051&start=20#p23867

They resemble the high-gain antenna as well.

The craft falls to 11.5Gm distance to Jupiter and acquires 4705m/s. 321kg hydrogen brake by 1378m/s in 40 days, leaving 2131kg on the orbit (3327m/s) of one moon of the Himalia group that has four or five members.
http://en.wikipedia.org/wiki/Himalia_group
The orbit is tilted by 27.5° so the wide launch window for this goal opens every six years.

A part of the upper Hohmann kick would better combine with the capture to benefit from some Oberth effect. I suppose the craft must arrive below the Himalia orbit and brake over half an ellipse or nearly 125 days while rising to the Himalia orbit. Someone else shall develop the theory or guidelines for crafts with weak accelerations.

The other small moons have unrelated orbit inclination; maybe some is accessible.

The probe sinks from the Amalia orbit (3327m/s tilted 27.5°) to a circular untilted 6890m/s orbit, of radius 4.29 times smaller, by a Hohmann transfer. 321kg hydrogen give in 40 days the 2035m/s upper kick that de-tilts (1582m/s) and brakes (1281m/s), leaving 1809kg on the 61d elliptic transfer. 255kg hydrogen in few kicks totalling 32d brake 1884m/s, leaving 1554kg on the circular 6890m/s orbit.

There, ten engines are thrown away (250kg), a 3905kg hydrogen tank (200kg), the truss that holds it (200kg) - though most of the truss could have been thrown right after the chemical propulsion. The probe continues at 904kg, braking slowly over many orbits as it consumes 15.5mg/s=1.34kg/d.

91kg hydrogen let dive by 1310m/s in 68 days and leave 813kg at 8200m/s around Jupiter, that is at Callisto.

158kg hydrogen let dive by 2683m/s in 118 days and leave 655kg at 10883m/s around Jupiter, that is at Ganymede.

135kg hydrogen let dive by 2862m/s in 101 days and leave 520kg at 13745m/s around Jupiter, that is at Europa. Here the probe might split a lander, throw away an engine or a tank... The following doesn't do it.

130kg hydrogen let dive by 3581m/s in 97 days and leave 390kg at 17326m/s around Jupiter, that is at Io.

The engines (50kg), the 514kg hydrogen tank (55kg), a truss around it (45kg) leave 240kg for the probe's frame, equipment, instruments.

Marc Schaefer, aka Enthalpy

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No probe has orbited Uranus; only Voyager 2 passed by in 1986, and knowledge improved slowly since then.
http://en.wikipedia.org/wiki/Uranus
http://en.wikipedia.org/wiki/Exploration_of_Uranus
http://en.wikipedia.org/wiki/Voyager_2

We won't wait one and a half decade for a Hohmann transfer. The Solar thermal rocket shall send the probe there in 6 years, which means 12858m/s relative to Earth and 12355m/s relative to Uranus, without the assistance of other planets.
[xls file joined]

An Atlas V v551 (or an Ariane V Me - something with a wide fairing) shall put 5589kg to 4300m/s above Earth's gravity. Four D=4.57m Solar engines add 8558m/s to the remaining 2806kg as they eject 2783kg hydrogen in 42 days. The Solar thermal stage is thrown away before braking: 155kg tank, 300kg truss, 120kg engines, leaving 2231kg.

A chemical rocket brakes by Oberth effect at the final orbit's periapsis:

  • Small radius 30Mm and 19173m/s
  • Big radius 600Mm and 959m/s
  • Period 5.34 days
  • Orientation as possible

The escape speed is 19646m/s at 30Mm, so the probe arrives there with 23208m/s and must lose 4035m/s.

The rocket burns 1282kg of 700:100 O2:H2 at 25 bar expanded to 84Pa in four D=0.8m nozzles to achieve 4699m/s=479s isp. Electric pumps take 14kW during 15min from a 28kg Li-polymer battery.
http://www.scienceforums.net/topic/73571-rocket-engine-with-electric-pumps/
The pumped engines weigh 40kg with driver, the tanks 40kg, their supporting truss 100kg, leaving 769kg for the probe's frame, instruments, equipment including the battery.

Few orbits are possible; this one looks useful. Landmarks:

Marc Schaefer, aka Enthalpy

 

========================================================================

 

Neptune resembles Uranus: similar mass and diameter, no orbiter ever, Voyager 2 passed by in 1989.
http://en.wikipedia.org/wiki/Neptune
http://en.wikipedia.org/wiki/Neptune#Exploration
http://en.wikipedia.org/wiki/Voyager_2
Sending an orbiter there is about the same, just farther.

An 8 year travel needs 14380m/s versus Earth and 15771m/s versus Neptune. The Solar thermal rocket starts again with 5589kg at 4300m/s above Earth's gravity, ejects 3106kg hydrogen in 47 days through four D=4.57m Solar engines to add 10080m/s, ending at 2483kg. 590kg of Solar thermal propulsion are thrown away.

The probe arrives with 1893kg at Neptune to brake chemically by Oberth effect to the final orbit:

  • Periapsis 28Mm, apoapsis 770Mm, period 7 days, orientation as possible.
  • Cf equator 24.8Mmn, Voyager 29.2Mm, rings 41-64Mm, Triton 355Mm, known moons 48Mm-far
  • At 28Mm, escape speed is 22091m/s, elliptic orbit 21700m/s

Rings http://en.wikipedia.org/wiki/Rings_of_Neptune
Moons http://en.wikipedia.org/wiki/Moons_of_Neptune
Magnetosphere http://en.wikipedia.org/wiki/Neptune#Magnetosphere

The probe arriving at 15771m/s dives to 27143m/s at 28Mm where it brakes by 5443m/s in 15min. This consumes 1299kg of O2 and H2 in the same 479s engine as at Uranus, leaving 594kg in orbit. 170kg for propulsion allow 386kg for frame, equipment, instruments.

===============

One fascinating option is to send sistership probes to Uranus and Neptune. Share the design, the production, the spare, the operation team, the science team - and launch both at the same epoch. Then, as teams are occupied by Uranus after 6 years travel, a 9.8 year travel to Neptune is more acceptable.

This takes 13230m/s from Earth, just 3% mass difference with the acceleration toward Uranus, and arriving with 13201m/s at Neptune takes a 4035m/s insertion kick at the described orbit - same kick as at Uranus, same 769kg available.

Marc Schaefer, aka Enthalpy

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How far from the Sun does diminishing radiation limit the use of your design?

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The scenario I propose for Jupiter brakes there using the Solar thermal rocket.

Same for the asteroids and the Jovian Troyans, which are return missions.

I didn't check for Uranus and suppose it's impossible at Neptune.

 

I know from previous scenarios that Mars is an easy target for the Solar thermal rocket.

Work is in progress for Saturn, I'll try a bang-and-whistles mission like at Jupiter, if possible.

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Hydrogen is the propellant that improves the exhaust speed over chemical reactions, but the Solar thermal engine accepts other propellants. Water at 2400K can give some 3000-4000m/s ejection speed, depending on the dissociation allowed by the chamber pressure - and if available in space, it enables big scenarios where the in-situ propellant needs no lengthy preparation.

Corrosion is a serious worry, hence the 2400K. If metals don't survive hot vapour, ceramics may: MgO and ZrO2, with 100K less? Tantalum hafnium carbide?

Imagine that we find main-belt comets of the proper size and clean enough
http://en.wikipedia.org/wiki/Main-belt_comet
they can refill a spacecraft's tank after a simple purification - faster and lighter than electrolysis.

A part of the icy object can fuel the Solar rocket to bring the rest (or a different asteroid) to a remote location.

  • To a Lunar orbit: too easy now with the Solar thermal engine! "Think big" means again: manned Mars mission.
  • To a Martian orbit or Mars' surface, for more ambitious uses.

Granting the engine some time enables a scale more pleasant than our present space tinkering. Take thirty 4.7m concentrators that fit in one launch: at mean 2 AU, each ejects 40kg/day vapour at 4030m/s, or together 2,200t in 5 years. As the travel from the main belt to Mars takes 4820m/s, it leaves 1000t water at Mars. That's a swimming pool of 2m*50m*10m. Enough to refill a rocket there to return to Earth, possibly after separation into hydrogen and oxygen. Or to grow vegetables?

Mars' gravitation is only 30 times ice's heat of fusion, so chunks of some proper size may aerobrake and reach the ground but cause limited damage.

Sixty 12m concentrators fit in one SLS fairing. These would bring in 10 years 26,000t water to Mars, or 5m*D80m: that's a pond. A duly megalomaniac inventor blink.png (salut Guy Pi) proposed to veer a comet off course and smash it on Mars for terraforming; still not quite the proper size, but it's a first step.

Marc Schaefer, aka Enthalpy

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One candidate to resist corrosion by vapour at 2400K is ZrO2 stabilized with Y2O3. For conductivity and if possible, zirconia should only cover a metal like tungsten or tantalum.

If 2400K are sustainable, expansion from 2000Pa to 1Pa brings isp=3591m/s=366s.

Main belt comets orbit rather at 3.1AU, some with a small inclination; only 6 bigger are known up to now, many more are expected. Spiralling from there to Mars takes (my mistake) 7210m/s, leaving 1/7.45 of the initial mass.

The lower temperature reduces the heat leaks, but scaling only by the isp, and at mean 2.11AU from Sun, thirty D=4.7m engines eject 18g/s=1558kg/day vapour, or 2846t over 5 years, which would leave 441t near Mars with little inert mass.

Sixty D=12m engines from an SLS fairing would deliver over 10 years 11500t water near Mars. The D=80m pond is 2.3m deep.

Several thousand tonnes water boiled in few dm3 must be very pure to avoid scaling. Do it as you can.

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Some uses and engines need water, others oxygen and hydrogen, still others just hydrogen. At Mars, the same thirty D=4.7m concentrators can power turbines to produce 137kW electricity if pessimistic
http://saposjoint.net/Forum/viewtopic.php?f=66&t=2051
of which 60% efficient electrolysis splits 0.29mol/s=450kg/day. The whole 440t take 2.7 years, resulting in 49.2t hydrogen.

A cryocooler to keep the propellants liquid is described in the same linked topic.

Whether water, only hydrogen, or as well oxygen are wanted, I find that transporting ice from the main belt by ejecting vapour takes fewer concentrators than making hydrogen at the main belt.

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At Mars or Earth, propellants brought from outer space are better kept near the planet than landed: the lighter crew or payload shall join the heavier propellants. An elliptic orbit, a high orbit or a Lagrange point are candidates, as is known.

Marc Schaefer, aka Enthalpy

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Usual honeycomb sandwiches with 0.3mm aluminium skins are heavier than desired for the concentrator. Space technology has probably lighter ones. Here I suggest a different process.

post-53915-0-40001100-1375028251_thumb.png

Nickel parts are produced down to 8µm thin for elastic couplings; here 50µm shall provide enough local stiffness, with features about 20mm wide for ~80mm periodicity, as an experience-backed feeling. A few additional parts shall provide the global stiffness; consider brazing.

The skin is deposited on a mould that gives the concentrating shape and preferably the optical smoothness. The reflecting layer may be deposited before. Then a thick layer is deposited temporarily on the skin and patterned to the hills and valleys; rocket cooling jackets use wax for that. The nickel stiffeners are deposited on the skin and the hills. The temporary layer is removed.

A rear skin could be added on the stiffeners using a new temporary thick layer to produce a complete sandwich. This can be lighter if the skins get thinner.

Once the several layers achieve enough stiffness, the reflector can be removed from the mould.

Other parts can be produced this way, for instance antennas, or Sunlight concentrators to produce electricity.

Other possibilities exist, for instance a sandwich of metal skins and thin balsa wood. The heat conductance suffices, but I like the all-nickel contruction for uniform thermal expansion.

The Astromesh antenna is big, unfoldable, and its weight would nearly fit, but it's not a precise and efficient optical reflector up to now.

Marc Schaefer, aka Enthalpy

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Atlas' capability on escape trajectories is less than I estimated, here a new attempt:
post-53915-0-35409200-1376166588.png

Mass at all steps of the mission can be multiplied by 0.78 or 0.78 - or take an Atlas V heavy.

 

Because the Esc-A stage weighs fat 4540kg, Ariane V underperforms Atlas V 551. The Esc-B is rumoured near shattering 5650kg after removing the equipment case, so Ariane V ME would but outperform Atlas V 551.

Even the specially-designed Esc-B weighs 200kg per ton of propellants and Ariane 6 shall use a similar stage - the competitors have 112kg/t at the antique Centaur and develop composite tanks. Time to wake up maybe?

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Europa is a big Jovian moon; prevalent models imagine beneath its thick ice crust a water ocean where wild scenarios put life.
http://en.wikipedia.org/wiki/Jupiter
http://en.wikipedia.org/wiki/Moons_of_Jupiter
http://en.wikipedia.org/wiki/Europa_(moon)
The Galileo probe passed by repeatedly; a special mission carrying recent instruments just to Europa has strong support.
http://en.wikipedia.org/wiki/Galileo_spacecraft
Europa's location (go to Jupiter, plunge to the 13739m/s circular orbit) challenges chemical rockets. A mission with fission reactor and ion thruster was abandoned. The Solar thermal engine achieves the performance naturally.

A Falcon-9 shall put 3414kg at 2945m/s above Earth's gravity, using an added escape stage:
http://saposjoint.net/Forum/viewtopic.php?f=66&t=2272&p=41436#p41436
after which nine D=4.572m Solar thermal engines add 5848m/s in eight days so 2132kg reach Jupiter in 33 months. Europa's orbit inclination opens two wide launch windows in 11.86 years.

A bigger rocket achieves more; so do slingshots by Venus, Venus and Earth, which smarter people can evaluate.

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If believing my spreadsheet, a capture sequence with weak braking should pass as low as possible by the celestial body, to match the periapsis of the capture orbit and the final orbit. The benefit is real even with 400µm/s2. It needs a pass but higher than the targeted periapsis, at least when braking always against the speed; pushing forward or downward at apoapsis may change the optimum.

 

EuropaWeakBrakeAtJupiter.zip

 

5.204AU to Sun let each engine consume 0.67kg/day and brake by 98mN, or together 414µm/s2 when arriving at Jupiter with 5643m/s, increasing as the craft lightens. According to my spreadsheet, they decelerate to 18535m/s at 712Mm from Jupiter's center in 97 days; 7.4 days more braking inject the craft on a 671Mm/13344Mm orbit with 120 days period. This uses 4020+306m/s. At the next periapsis, a long kick of 306+306m/s brings the period to 57 days and the apoapsis to 4287Mm. Combined, they consume 4938m/s or 699kg hydrogen, leaving 1433kg on the 671Mm/4287Mm Jupiter orbit.

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To target Europa and its circular 671Mm Jupiter orbit, multiple "short" kicks at periapsis cost 4934m/s or 470kg hydrogen, leaving 963kg. If pushing over 20% of the orbit's duration, it takes 390 days, not quite pleasant.

Smarter people would probably brake by many successive slingshots at Europa, whose period is only 3.5512 days, or at the heavier Ganymede. This should save propellants AND time.

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Capture by Europa isn't obvious, because weak braking leaves a high apoapsis, but Europa's influence extends only 15Mm on the Jupiter-Europa line. The apoapsis takes time to lower, and meanwhile Europa moves much around Jupiter, so the still high apoapsis can become aligned with Jupiter, bye-bye.

We can increase the engines' thrust so the apoapsis lowers quickly enough, but my estimates tell that the lower engine performance makes this option less good.

The better option I've seen puts the capture apoapsis over a pole, where Jupiter won't eject the craft. This permits the polar orbit preferred for an exploration probe. The small needed North-South speed before capture, like 100m/s, results from a slightly tilted 671Mm Jupiter orbit. The approach would leave no relative East-West speed, hence waive its favourable interaction with the capture, but this looks globally favourable.

EuropaWeakBrake.zip

 

Braking 4.4 days before to 0.1day after the pass at 1.83Mm and 1821m/s injects on a 1.8Mm/29Mm polar Europa orbit. Few weeks suffice to the circular 1.8Mm (1800km) Europa polar orbit - little over Europa's 1561km radius. The sequence costs 855m/s or 64kg hydrogen, leaving 899kg.

The engines weigh around 270kg; some could be jettisoned while lowering the Jupiter orbit. The tank for 2515kg hydrogen can be split; as a single D=4.1m suspended balloon, it could weigh 115kg and its surrounding truss ~120kg.

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899kg, obtained from a medium launcher, is a nice mass for an orbiter with cameras and deep-view radar.

Slingshots at Venus, Earth and Europa would increase that, enabling additional lander and diver.

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Encelade is also a Moon with an ice crust and supposedly liquid water beneath, this one at Saturn. A similar mission there looks feasible, with comparable speed requirements, but is more difficult due to the fainter Sunlight. Concentrators lighter than 1kg/m2 would help even more.

Marc Schaefer, aka Enthalpy

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